A gas turbine engine typically includes a turbomachinery core having a high pressure compressor, combustor, and high pressure turbine in serial flow relationship. The core is operable in a known manner to generate a primary gas flow. The high pressure compressor includes annular arrays (“rows”) of stationary vanes that direct air entering the engine into downstream, rotating blades of the compressor. Collectively one row of compressor vanes and one row of compressor blades make up a “stage” of the compressor. Similarly, the high pressure turbine includes annular rows of stationary nozzle vanes that direct the gases exiting the combustor into downstream, rotating blades of the turbine. Collectively one row of nozzle vanes and one row of turbine blades make up a “stage” of the turbine. Typically, both the compressor and turbine include a plurality of successive stages.
Gas turbine engines, particularly aircraft engines, require a high degree of periodic maintenance. For example, periodic maintenance is often scheduled to allow internal components of the engine to be inspected for defects and subsequently repaired. Unfortunately, many conventional repair methods used for aircraft engines require that the engine be removed from the body of the aircraft and subsequently partially or fully disassembled. As such, these repair methods result in a significant increase in both the time and the costs associated with repairing internal engine components.
Gas turbine engines include various rotors in the typical form of bladed disks. Each rotor disk is specifically configured with a radially outer rim from which extends a row of blades. An axially thinner web extends radially inwardly from the rim and terminates in an axially thicker hub having a central bore therein.
A particular advantage of the bladed disk construction is that the integral disk may be smaller since no dovetails are used, and the blades are integrally formed around the disk rim. However, this construction increases repair difficulty since the blades are not readily individually removable from the disk. Minor repairs of the blade may be made in the bladed disk, but major repair thereof requires removal by cutting of corresponding portions of damaged blades or their complete removal, with the substitution thereof being made by welding or other metallurgical bonding process for achieving the original strength of the bladed disk.
An additional difficulty in the manufacture of the bladed disk is balancing thereof. All rotor components in a gas turbine engine must be suitably statically and dynamically balanced for minimizing rotary imbalance loads during operation for reducing vibration. The dovetail disk construction permits the rotor to be initially balanced during manufacture, with the individual blades being separately manufactured and matched in position on the disk for minimizing the resulting imbalance of the assembly thereof.
As such, a need exists for a method of in situ balancing of an internal rotating component, particularly a rotating disk, of a gas turbine engine.